Pulsed thrust velocity control of a projectile



y 1955 R. F. MAY ETAL 3,184,182

PULSED THRUST VELOCITY CONTROL OF A PROJECTILEI Filed Jan. 18. 1960 3 Sheets-Sheet l Q, e v Q 2 Q3 $1 DIRECTION OF RIGHTING 0 Q MoMENT CAUSED BY NoRMAL FORCE TENDING To REDUCE ANGLE OF ATTACK ANGLE OF/ ATTACK I DIRECTION DIRECTION OF TRAVEL CENTER OF GRAVITY OF MISSILE AERODYNAMIC CENTER OF PRESSURE. THROUGH WHICH NORMAL FORCE ACTS DIRECTION OF ACCELERATION DUE TO GRAVITY DIRECTION OF ACCELERATION DUE TO DRAG DIRECTION OF ACCELERATION DUE TO TI-IRUST SCHEMATIC REPRESENTATION OF THE FORCES ON A MISSILE E-JH ELLIPTICAL PATH (zERo INERTIAL FORCES) LLI O D SQQ TRAJECTORY OF I MISSILE UNDER THE I- INFLUENCE OF j UNCoMPENsATED DRAG BOOST PHASE RANG E I REP RESENTATION OF MISSILE TRAJECTORIES INVENTORS. F4267; F/Yd mva 3 E- a Da A M May 18, 1965 R. F. MAY ETAL 3,184,182

PULSED THRUST VELOCITY CONTROL OF A PROJECTILE Filed Jan. 18. 1960 AVERAGE DRAG /fTHRusT PULSES 3 Sheets-Sheet 2 INVENTORS. Wa e?? P N E 5 David. L. Dari.

United States Patent 3,184,182 PULSED THRUST VELOCITY CONTROL 0F A PROJECTEE Robert F. May, Birmingham, and David L. Dart, Pontiac, Mich, assignors to Chrysler Corporation, Highland lark, Mich, a corporation of Delaware Filed Jan. 18, 1960, Ser. No. 3,139 Claims. (61. 24414) This invention relates to a method and the means for eliminating target errors accumulated by a missile in the post boost phase of its operation. The usual sources of these errors are unpredictable atmospheric conditions including winds, and known missile parameters such as missile geometry, weight and drag characteristics.

This invention assumes that the boost phase of the missiles flight has been accomplished without errors, that is, that the missile, by means not encompassed by this invention, has had imparted to it a velocity, known in both magnitude and direction. This velocity is such, in both magnitude and direction, that if no further inertial forces are applied, the resulting elliptical trajectory strikes the earth at the exact position of a desired target. If this elliptical path can be maintained, the missiles impact point is completely determined by the initial condition of velocity, both magnitude and direction, and is completely independent of post boost flight conditions of meteorological conditions, and missile flight parameters. The present invention provides the method and means for achieving this post boost phase elliptical path.

It is a principal object, therefore, of this invention to counteract the inertial forces caused by varying atmospheric and other non-readily measurable factors acting along the longitudinal axis of an aerodynamically stable missile during its post boost flight so that the missile will follow a substantially elliptical path to the target.

The accompanying drawings, graphs, and sketches are designed to facilitate a more rapid and complete understanding of applicants invention and comprise:

FIGURE 1 representing the principal forces acting on aerodynamically stable missiles or projectiles;

FIGURE 2 representing the effects of drag on the desired elliptical trajectories of missiles or projectiles;

FIGURE 3 representing the graph of thrust and drag versus time in a missiles flight through the atmosphere;

FIGURE 4 representing graphs of the missiles physical reactions to the inertial forces acting along its longitudinal axis during its post boost flight, and the conversion of these reactions to control data; and

FIGURE 5 representing a schematic view of the pulse and control systems of the present invention.

The inertial or non-gravitational forces acting upon a missile may be broken up into the two general components of (1) normal forces which act perpendicularly to the missiles longitudinal axis and, (2) those forces acting along the missiles longitudinal axis. The normal force is a result of angle of attack, that is, the angle between the missiles longitudinal axis and the direction of the relative wind (see FIGURE 1). The missile to which the present invention applies, is of the aerodynamically stable type that zeroes out normal force, except for transient conditions, by automatically aligning its longitudinal axis with the relative wind. Other forces normal to the body axis may result from a misalignment of both,

' or either, the thrust or aerodynamic axis of the missile.

For purposes of this invention, these unsymmetrical forces are rotated around the direction of flight by imparting a roll to the missile. Their eflect on the flight path is thus averaged to zero. Thus the effects of all inertial forces except those along the body longitudinal axis are eliminated.

The direct techniques for nulling the effects of nongravitational components of acceleration along the body axis would be to vary either the thrust of a propulsion system or the amount of additional drag (by controlled drag surfaces), but these methods involve restrictive practical limits and relatively low efiiciency. Several factors which contribute to the undesirability of these techniques are as follows: A missile travels through air of a density which varies significantly with altitude above the earth; also, it travels at different speeds, dependent upon the flight range desired; and drag variations during one flight and between flights of different ranges may be considerable with ratios of :1 or :1 total variations characteristic for short-range systems.

In the present invention, the effect of drag forces tending to decelerate the missile and causing it to depart from the desired elliptical path (FIGURE 2) are neutralized by the counteracting forces of a pulsed thrust engine. The pulses are commanded by a control system whose only sensory component consists of an accelerometer rigidly attached to the missile structure with its sensitive axis coincident with the missiles longitudinal axis. This accelerometer will be sensitive to all non-gravitational accelerations along the latter axis. The output of the pulsing thrust unit may be of a constant thrust level for each pulse, which level is set near the maximum drag anticipated for the missile over its complete spectrum of ranges. The pulses will be so timed and of such a duration to cancel the decelerating effects of drag. Thus, in a high drag environment there will be more engine pulses per unit time than will be required in a low drag environment. FIGURE 3 is a plot of drag force for an elliptical flight path versus flight time superimposed upon which is a representation of the engine thrust pulses. During the portion of the missiles flight through the more dense parts of the atmosphere which are. represented by the initial and terminal portions of FIGURE 3, the drag on the missile is high and a correspondingly high frequency of pulses is required to neutralize it. However, in the middle part of the flight when relatively high altitudes are attained, the atmospheric drag is slight and the thrust pulse intervals are, therefore, stretched out. The drag deceleration is affected by winds, atmospheric conditions, and missile geometry, and weight. As these parameters change from the standard predicted values, so will the frequency of thrust pulsations change from the predicted value to compensate.

The pulse engine is controlled by information derivable from the aforementioned longitudinal accelerometer. Also on board the missile are means for computing the first and second integrals of the measured acceleration. The accelerometer may be of the integrating type which gives directly the first integral of acceleration as output, which is then integrated once for the second integral of acceleration, or it may be the pendulous mass type in which case two integraters are employed. These control elements are diagrammed schematically in FIGURE 5. The control system operates valves which govern the fuel supply to the engine. The fuel and its oxidizer form a hypergolic mixture so that no ignition system is required. Examples of such mixtures are methanol and hydrogen peroxide, and aniline and nitric acid. The engine can be actuated to give constant or varying impulse by controlling thrust level and/ or pulse width. Thrust level and/ or pulse width can be changed either by control system criteria, or as timed functions depending on predicted trajectory conditions. The operation of the system is relatively insensitive to variations in total impulse per pulse.

The following method of control of the pulsing engine may be understood by comparing the pulsing missile with a hypothetical reference point that travels the desired elliptical path with inertial forces continuously zero. An accelerometer on board this hypothetical missile would read identically zero at all times and the first and second integrals would also be zero. The pulsating missile then with all normal forces averaged out would be aware of acceleration, velocity and displacement through its control intelligence, all with respect to the reference point. FIGURE 4 shows the missiles inertial response over one cycle to a control scheme which commands a pulse whenever the first integral, which represents missile velocity with respect to the reference point, is zero or less, and, concurrently, when the second integral, which represents missile displacement with respect to the reference point is zero or less. Stated in symbolic form, a pulse is commanded when .7 jAdt and jjAdt 0 where t is time, and where A is the acceleration as measured by the body fixed longitu dinal accelerometer. As'may be seen in FIGURE 4, when the first pulse is commanded, the missile accelerates with respect to the reference point-and therefore begins to travel faster and pulls ahead of the reference point. When the pulse is terminated, drag decelerates the missile and its velocity decreases, but since it is still positive, the displacement increases with respect to the reference point. Eventually the action of drag reduces the velocity indication to zero at which point the maximum diplacement from the reference point occurs. The velocity then becomes negative and the displacement begins to decrease. When the displacement reaches zero, the control conditions of the first and second integrals of acceleration being equal to or less than zero are satisfied so another pulse is commanded. This process repeats so that the average velocity of the pulsing missile during its post boost flight is zero with respect to the reference point and therefore must travel the same path, all inertial forces other than those along the longitudinal axis being zero.

Referring to FIGURE 5 of the drawing, a missile is shown having fins 12 thereon to make the missile aerodynamically stable with a thrust motor 14 of the type which can be inermittently operated to produce thrust pulsations of either constant or varying levels. As stated above, this motor preferably is of the type which may use hypergolic fuel and oxidizers so that practically instantaneous thrust pulsations may be had upon the contact of the fuel with the oxidizer within the motor. A tank 16 contains the hypergolic fuel and oxidizer in separate compartments and intermittently feeds them to the motor 14 by a valve 1-8 which keeps these fuel constituents unmixed until they combine within the motor 14. A solenoid valve 20 is provided to open and close the valve 18 upon receiving an electrical signal from a power source 22.

The electrical circuit which connects power source 22 to the valve 20 is provided with a coincidence gating means comprising two switches 24 and 26 arranged in series. Switch 24 is connected to a rectifier mechanism 28 which allows only negative electrical signals to be transmitted to a power element 29 such as a solenoid to cause switch 24 to close. The switches of FIGURE 5 are diagrammatic and do not necessaritly constitute the actual type of equipment to be used for this job. The integrated missile velocity change electrical signals received in rectifier 28 are put out by an integrator 30 which delivers the first integral of missile velocity change as measured by an accelerometer 32 mounted with its sensitive axis coincident with the longitudinal axis XX of the missile.

Another integrator 34 integrates the output of integrator 30 and transmits the resulting missile displacement signals to another rectifier 36 which allows only negative displacement signals to be transmitted to power unit 37 and cause switch 26 to close. When the condition represented by the equations fAdt 0 and ffAdt 0 exists, switches 24 and 26 are both closed and the motor 14 gives a thrust pulsation of predetermined duration.

The present invention may also be carried out by setting V=fAdt and S=jfAdt equal to a range of values the breadth of which will be determined by trajectory requirements. However, the preferred embodimentof the invention sets these control equations concurrently equal to or less than zero.

We claim:

1. A velocity sustainer system for counteracting atmospheric drag on a projectile during its post boost phase and thereby allowing the projectile to fly a substantially elliptical locus comprising accelerometer means rigidly mounted on the projectile with its sensitive axis lying along the longitudinal axis of the projectile for measuring the inertial acceleration thereof due to engine thrust and atmospheric force along the longitudinal axis of the projectile, an electrically operable propulsion means for directing an accelerating force along the longitudinal axis of the projectile, first integrating means electrically coupled'to said accelerometer and operable to provide an electrical output representative of velocity relative to a reference point on the locus, second integrating means electrically coupled to the output of said first integrating means and operable to provide an electrical output representative of displacement from the reference point on the locus, first and second switching means coupled in series and operatively connected to and controlling by their simultaneous operation the operation of said propulsion means, said first switching means operable responsive to a predetermined voltage output from said first integrating means and said second switching means operable responsive to a predetermined voltage output from said second integrating means.

2. A velocity sustaincr system for counteracting atmospheric drag on a projectile during its post boost phase and thereby allowing the projectile to fly a substantially elliptical locus comprising accelerometer means rigidly mounted on the projectile with its sensitive axis lying along the longitudinal axis of the projectile, an electrically operable propulsion means for directing accelerating forces of predetermined duration along the longitudinal axis of the projectile, first integrating means electrically coupled to said accelerometer and operable to provide an electrical output representative of velocity relative to a reference point on the locus, second integrating means electrically coupled to the output of said first integrating means and operable to provide an electrical output representative of displacement from the reference point on the locus, first and second switching means coupled in series and operatively connected to and controlling the operation of said propulsion means, and means for actuating said first and second switching means, said actuating means electrically coupled to and operable responsive to a negative output from said first and second integrating means.

3. The combination as set forth in claim 1 wherein said electrically operable propulsion means comprises an electrically valved thrust motor of the bipropellant type.

4. The combination as set forth in claim 2 wherein said actuating means comprises a first solenoid operatively connected to and controlling said first switching means and a second solenoid operatively connected to and controlling said second switching means, and wherein a first rectifier is connected between said first integrating means and said first solenoid and a second rectifier is connected between said second integrating means and said second solenoid, each of said rectifiers poled to transmit signals of negative polarity.

5. The combination as set forth in claim 2 wherein said electrically operable propulsion means comprises an electrically valved thrust motor of the bipropellant type.

References Cited by the Examiner UNITED STATES PATENTS 2,613,071 10/52 Hansel 324-70 2,932,467 4/60 Scorgie 24414 3,073,550 1/63 Young 244-14 SAMUEL FEINBERG, Primary Examiner.

CHESTER L. JUSTUS, KATHLEEN CLAFFY,

Examiners. 

1. A VELOCITY SUSTAINER SYSTEM FOR COUNTERACTING ATMOSPHERIC DRAG ON A PROJECTILE DURING ITS POST BOOST PHASE AND THEREBY ALLOWING THE PROJECTILE TO FLY A SUBSTANTIALLY ELLIPTICAL LOCUS COMPRISING ACCELEROMETER MEANS RIGIDLY MOUNTED ON THE PROJECTILE WITH ITS SENSITIVE AXIS LYING ALONG THE LONGITUDINAL AXIS OF THE PROJECTILE FOR MEASURING THE INERTIAL ACCELERATION THEREOF DUE TO ENGINE THRUST AND ATMOSPHERIC FORCE ALONG THE LONGITUDINAL AXIS OF THE PROJECTILE, AN ELECTRICALLY OPERABLE PROPULSION MEANS FOR DIRECTING AN ACCELERATING FORCE ALONG THE LONGITUDINAL AXIS OF THE PROJECTILE, FIRST INTERGRATING MEANS ELECTRICALLY COUPLED TO SAID ACCELEROMETER AND OPERABLE TO PROVIDE AN ELECTRICAL OUTPUT REPRESENTATIVE OF VELOCITY RELATIVE TO A REFERENCE POINT IN THE LOCUS, SECOND INTERGRATING MEANS ELECTRICALLY COUPLED TO THE OUTPUT OF SAID FIRST INTEGRATING MEANS AND OPERABLE TO PROVIDE AN ELECTRICAL OUTPUT REPRESENTATIVE OF DISPLACEMENT FROM THE REFERENCE POINT ON THE LOCUS, FIRST AND SECOND SWITCHING MEANS COUPLED IN SERIES AND OPERATIVELY CONNECTED TO AND CONTROLLING BY THEIR SIMULTANEOUS OPERATION AND OPERATION OF SAID PROPULSION MEANS, SAID FIRST SWITCHING MEANS OPERABLE RESPONSIVE TO A PREDETERMINED VOLTAGE OUTPUT FROM SAID FIRST INTEGRATING MEANS AND SAID SECOND SWITCHING MEANS OPERABLE RESPONSIVE TO A PREDETERMINED VOLTAGE OUTPUT FROM SAID SECOND INTEGRATING MEANS. 